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Abstract:

Disclosed herein, in certain embodiments, is a method of altering the
stability of unstable space debris. In some embodiments, the method
further comprises changing the orbit of the unstable space debris.

Claims:

1-14. (canceled)

15. A satellite for stabilizing unstable space debris, comprising (a)
routine systems and subsystems for the operation of the satellite, and
(b) a means for generating and directing a gas plume sufficient to dampen
the rotational momentum about one or more of the axes of rotation of the
unstable space debris.

16. The satellite of claim 15, further comprising an active control
system for maintaining the position and attitude of the satellite during
proximity operations and while using the gas impingement system to
stabilize or change the orbital parameters of the unstable space debris.

17. The satellite of claim 15, further comprising a means for analyzing
the motion of the unstable space debris.

18. The satellite of claim 15, further comprising a laser tracking
system, a radar (or other radio frequency) tracking system, an optical
tracking system, or a combination thereof.

19. The satellite of claim 15, further comprising a laser or radar
tracking system and an optical tracking system.

20. The satellite of claim 15, further comprising a means for calculating
the strength of the pulse of the gas plume, and the number, duration, and
timing of the pulses of the gas plumes to be applied to the unstable
space debris.

21. The satellite of claim 15, further comprising a means of being
serviced and refueled so that it can stabilize, capture or change the
orbits of multiple pieces of space debris without have to be de-orbited
itself or without have to have new satellites launched into space.

22. (canceled)

23. (canceled)

Description:

CROSS-REFERENCE

[0001] This application claims priority from U.S. Provisional Patent
Application 61/264,386, filed Nov. 25, 2009, which is incorporated herein
by reference in its entirety.

BACKGROUND OF THE INVENTION

[0002] There are over 20,000 pieces of man-made (artificially introduced)
space debris currently being tracked in orbit around the Earth. Some
space debris is found in Low Earth Orbit (LEO) (e.g., at altitudes
between 200 km and 2,000 km). Some space debris is found in Medium Earth
Orbit (MEO) (e.g., at altitudes between 2,000 km and 35,586 km). Some
space debris is found in Geosynchronous Earth Orbit (GEO) (e.g., at
altitude of 35,786 km±200 km). Some space debris is found at altitudes
in excess of the GEO belt. Much of the space debris (approximately 40%)
is concentrated in stable circular or elliptical orbits between 200 km
and 2,000 km.

SUMMARY OF THE INVENTION

[0003] Space debris is a growing issue threatening the ability to safely
operate spacecraft in Earth Orbit. Because space debris is not under
active positive control, it represents a collision hazard to spacecraft.
Currently, it is the spacecraft that is acted upon (e.g., maneuvered to
avoid a collision). However, maneuvering to avoid debris is costly. It
artificially decreases the useful life of the spacecraft (e.g.,
satellites or payloads) by expending fuel that is intended for other
mission essential events.

[0004] Alternatively, the unstable space debris can be acted upon--it can
be removed from the orbital path of a spacecraft. While the capability to
rendezvous with space objects and conduct proximity operations in the
vicinity of those objects has been demonstrated, there is currently no
demonstrated means of stabilizing unstable space debris so that (a) the
unstable space debris may be captured for de-orbit or (b) the orbital
parameters of the debris may be changed to either maneuver the debris
into a safe orbit or for de-orbit. Safe techniques to stabilize space
debris in a zero gravity and zero pressure environment is one of the
technical challenges that has yet to be addressed to by the space
community. Thus, there is a need for a method of stabilizing unstable
space debris.

[0005] Disclosed herein, in certain embodiments, is a method of
stabilizing unstable space debris, comprising: applying force to the
unstable space debris at target points on the unstable space debris,
generating stabilized space debris; wherein the force is generated by
pneumatic impingement of the unstable space debris with a gas plume
applied by an adjacent satellite; and wherein the force is sufficient to
generate a torque on the unstable space debris that will dampen the
rotational momentum about one or more of the axes of rotation of the
unstable space debris. In some embodiments, the amount of force being
applied to the unstable space debris is a function of the motion of the
unstable space debris, the state vector for the Center of Mass of the
unstable space debris, the mutually orthogonal axes of rotation of the
unstable space debris, the Moments of Inertia of the unstable space
debris, the rotational momentum of the unstable space debris, the target
points of the unstable space debris, or any combination thereof. In some
embodiments, the amount of force being applied does not damage the target
points. In some embodiments, the target points are located on, or near
to, each of the three mutually orthogonal axes of rotation centered at
the Center of Mass. In some embodiments, each target point is (a) located
on, or near to, each of the three mutually orthogonal axes of rotation
centered at the Center of Mass, and (b) structurally rigid enough to
absorb the force without being compromised. In some embodiments, the
target point(s) on the unstable space debris are a function of the Center
of Mass, the direction and magnitude of the velocity vector of the Center
of Mass, the moments of inertia and the rotational momentum about the
mutually orthogonal axes of rotation of the body of the unstable space
debris, or any combination thereof. In some embodiments, the number of
pulses of the gas plume required to stabilize one of the three mutually
orthogonal axes of rotation centered at the Center of Mass is independent
of the number of pulses of gas required to stabilize the other two axes
of rotation. In some embodiments, the gas plume comprises a gas selected
from: nitrogen gas; xenon gas; argon gas; neon gas; high velocity
residual affluent from chemical combustion of an oxidizer and a
propellant; high velocity residual affluent from the exothermal chemical
decomposition of a monopropellant on a catalyst; hydrogen gas; helium
gas; or a combination thereof. In some embodiments, the gas plume issues
from a nozzle selected from a divergent nozzle, a convergent nozzle, and
a collimated nozzle. In some embodiments, the gas plume issues from a
nozzle adjacent to at least one target point. In some embodiments, the
gas plume issues from a nozzle adjacent to one of the axes of rotation of
the unstable space debris. In some embodiments, the gas plume issues from
a nozzle found on a mechanically deployable arm. In some embodiments, the
method further comprises capturing the stabilized space debris. In some
embodiments, the method further comprises changing the orbital parameters
of the stabilized space debris.

[0006] Disclosed herein, in certain embodiments, is a satellite for
stabilizing unstable space debris, comprising (a) routine systems and
subsystems for the operation of the satellite, and (b) a means for
generating and directing a gas plume sufficient to dampen the rotational
momentum about one or more of the axes of rotation of the unstable space
debris. In some embodiments, the satellite further comprises an active
control system for maintaining the position and attitude of the satellite
during proximity operations and while using the gas impingement system to
stabilize or change the orbital parameters of the unstable space debris.
In some embodiments, the satellite further comprises a means for
analyzing the motion of the unstable space debris. In some embodiments,
the satellite further comprises a laser tracking system, a radar (or
other radio frequency) tracking system, an optical tracking system, or a
combination thereof. In some embodiments, the satellite further comprises
a laser or radar tracking system and an optical tracking system. In some
embodiments, the satellite further comprises a means for calculating the
strength of the pulse of the gas plume, and the number, duration, and
timing of the pulses of the gas plumes to be applied to the unstable
space debris. In some embodiments, the satellite further comprises a
means of being serviced and refueled so that it can stabilize, capture or
change the orbits of multiple pieces of space debris without have to be
de-orbited itself or without have to have new satellites launched into
space.

[0007] Disclosed herein, in certain embodiments, are methods of
stabilizing unstable space debris, comprising: applying a force to one or
more target point(s) located on mutually orthogonal axes that are
centered on the Center of Mass (CM) of the unstable space debris, wherein
the force(s) produce a torque on the unstable space debris that is
sufficient to dampen the rotational momentum about one or more of the
axes of rotation of the unstable space debris, and wherein the force is
generated and applied by an adjacent satellite. In some embodiments,
applying a force comprises impingement of the unstable space debris with
a gas plume generated by the adjacent satellite. In some embodiments, one
pulse of a gas plume impinges on the unstable space debris. In some
embodiments, multiple pulses of a gas plume impinge on the unstable space
debris. In some embodiments, the number of pulses of a gas plume required
to stabilize an axis of motion on the unstable space debris is
independent of the number of pulses of a gas plume required to dampen the
moments of inertia about the other axes of motion of the unstable space
debris. In some embodiments, the gas plume consists of: nitrogen gas;
xenon gas; argon gas; neon gas; gaseous ammonia; freon gas; high pressure
residual affluent from chemical combustion of an oxidizer and a
propellant; high pressure residual affluent from the chemical reaction
between a monopropellant (e.g., hydrazine, monomethylhydrazine, a variant
thereof, or hydrogen peroxide), and a catalyst; hydrogen gas; helium gas;
ionized cesium; ionized mercury; plasmas generated from compounds such as
teflon; or a combination thereof. In some embodiments, the gas plume is
generated by compressing the gas. In some embodiments, the gas plume is
generated by combustion of an oxidizer and a propellant. In some
embodiments, the gas plume is generated by the exothermic chemical
decomposition of a monopropellant reacting with a catalyst. In some
embodiments, the gas plume is generated by electrothermal, electrostatic,
or electromagnetic acceleration of one or more propellants. In some
embodiments, target point(s) on the unstable space debris are determined
by computer analysis, human analysis, or a combination thereof. In some
embodiments, target point(s) on the unstable space debris are a function
of (a) the Center of Mass, (b) the direction and magnitude of the
rotation about the Center of Mass of the body of the unstable space
debris and (c) the direction and magnitude of the velocity vector of the
Center of Mass of the body of the unstable space debris. In some
embodiments, target point(s) on the unstable space debris are a function
of the three mutually orthogonal axes of motion centered at the Center of
Mass. In some embodiments, target point(s) on the unstable space debris
have sufficient structurally rigidity to absorb the applied force without
being compromised (i.e., damaged). In some embodiments, target point(s)
on the unstable space debris are (a) on, or near, one or more of the
three mutually orthogonal axes of motion, and (b) have sufficient
structurally rigidity to absorb the applied force without being
compromised (i.e., damaged). In some embodiments, the amount of force
applied to the unstable space debris is a function of (a) the motion of
the debris, (b) the state vector for the Center of Mass of the unstable
space debris, (c) the mutually orthogonal axes, (d) the Moments of
Inertia, (e) rotational momentum, (f) the distance of the target points
from the Center of Mass, or (g) any combination thereof. In some
embodiments, the force applied to the unstable space debris does not
exceed the structural limitations of the target points. In some
embodiments, the force is generated and applied by a single satellite
with multiple mechanically articulated arms. In some embodiments, the
force is generated and applied by multiple adjacent satellites. In some
embodiments, the force is generated and applied by one adjacent satellite
per axis of rotation. In some embodiments, the force is generated and
applied by multiple satellites per axis of rotation. In some embodiments,
the method further comprises altering the orbital path of the unstable
space debris. In some embodiments, the method further comprises capturing
the unstable space debris for de-orbit. In some embodiments, the method
further comprises changing the orbital parameters of the unstable space
debris.

[0008] Disclosed herein, in certain embodiments, are satellites for
stabilizing unstable space debris, comprising: (a) satellite bus with
standard subsystems and interfaces, (b) a means for generating and
projecting a force sufficient to dampen the rotational momentum about one
or more of the axes of rotation of the unstable space debris, (c) one or
more external sensors (radar, laser radar (LIDAR), optical, or imaging
sensors), and (d) and an electronic system designed to analyze the data
from the sensors and develop a stabilization plan which includes the
force, duration, number, direction and magnitude of the pneumatic (gas)
jets (plumes). In some embodiments, the satellite further comprises an
active control system for maintaining the position of the satellite. In
some embodiments, the satellite further comprises a reaction control
system (RCS), control moment gyroscopes (CMG), magnetic torque converters
for attitude control, or a combination thereof. In some embodiments, the
satellite further comprises a means for remotely scanning and analyzing
the motion of the unstable space debris. In some embodiments, the
satellite further comprises a laser tracking system. In some embodiments,
the satellite further comprises a radar tracking system. In some
embodiments, the satellite further comprises an optical tracking system.
In some embodiments, the satellite further comprises a means for
calculating the amount of force to be applied to the unstable space
debris (the number, duration and timing of pneumatic (gas) pulses
(plumes) to be projected towards and impinge on the target points on the
unstable space debris) or a combination thereof. In some embodiments, the
satellite further comprises an on-board computer module. In some
embodiments, the satellite further comprises a de-orbit module for
attachment to the stabilized space debris. In some embodiments, the
de-orbit module comprises a rocket motor and fuel with sufficient thrust
to put the space debris into a reentry path and a guidance, navigation
and control system. In some embodiments, the satellite further comprises
a means for collecting the unstable space debris. In some embodiments,
the satellite further comprises a deployable robotic arm. In some
embodiments, the satellite further comprises a container for capturing
stabilized space debris. In some embodiments, the satellite further
comprises a means for storing the captured space debris. In some
embodiments, the satellite further comprises a container for storing the
captured space debris. In some embodiments, the satellite is serviceable.
In some embodiments, the satellite is refuelable.

[0009] Disclosed herein, in certain embodiments, is space debris that is
stabilized by a method disclosed herein.

[0010] Disclosed herein, in certain embodiments, is space debris that is
captured by a method disclosed herein.

[0011] Disclosed herein, in certain embodiments, is space debris that is
de-orbited by a method disclosed herein.

[0012] Disclosed herein, in certain embodiments, is space debris that is
stabilized by use of the satellite disclosed herein.

BRIEF DESCRIPTION OF THE DRAWINGS

[0013] The novel features of the invention are set forth with
particularity in the appended claims. A better understanding of the
features and advantages of the present invention will be obtained by
reference to the following detailed description that sets forth
illustrative embodiments, in which the principles of the invention are
utilized, and the accompanying drawings of which:

[0014] FIG. 1 illustrates the coordinate system and axes of motion of
unstable space debris.

[0022] Space debris includes inoperative satellites and payloads, expended
rocket stages which are not de-orbited, and the residual debris from of
satellites and payloads which catastrophically fail. While functioning
normally, satellites and payloads are actively or passively controlled by
internal control systems. At the "end of life" (EOL) of these systems,
those satellites and payloads that are not de-orbited are uncontrolled
and therefore may become unstable. Additionally, satellites and payloads
may unexpectedly fail before the designed EOL at which time they may
become uncontrollable and unstable. As a result they may begin to tumble
or rotate in one or more axes of motion (becoming unstable). Expended
rocket stages which are not de-orbited, and residual debris from of
satellites and payloads which catastrophically fail, may have no control
system and their rotational motion may also be unstable.

[0023] Space debris is a growing issue threatening the ability to safely
operate spacecraft in Earth orbit. Because space debris is not under
active positive control, it represents a collision hazard to other
spacecraft. Currently, it is the operative spacecraft that are acted upon
(e.g., maneuvered to avoid a collision). However, maneuvering to avoid
debris is costly. It artificially decreases the useful life of
operational satellites (and the associated payloads) by expending fuel
that is intended for other mission essential events.

[0024] Alternatively, the unstable space debris can be acted upon--it can
be removed from the orbital path of a spacecraft. While the capability to
rendezvous with space objects and conduct proximity operations in the
vicinity of those objects has been demonstrated, there is currently no
demonstrated means of stabilizing space debris with unstable rotation
motion (hereinafter, "unstable space debris") so that (a) the space
debris may be captured for de-orbit or (b) the orbital parameters of the
space debris may be changed. Safe techniques to stabilize unstable space
debris in a zero gravity and zero pressure environment is one of the
technical challenges that has yet to be demonstrated. Thus, there is a
need for methods of stabilizing unstable space debris.

Certain Terminology

[0025] As used herein, "satellite" means any object which has been placed
into orbit by human endeavor. In some embodiments, the satellite is
capable of autonomous control. In some embodiments, the satellite is
controlled by a ground-based operator.

[0026] As used herein, "space debris" means inoperative man-made objects
found in space and natural objects found in space. In some embodiments,
space debris is the residue of launch (inoperative boosters or debris
that is released from a rocket body during powered operation); the
residue of staging (explosive bolt residue or other hardware that is
deployed or activated during staging); satellites or payloads that
unexpectedly cease to operate or are uncontrollable but otherwise are
intact; satellites or payloads that catastrophically fail; satellites or
payloads that break up by causes other than catastrophic failure;
satellites or payloads that are damaged or break up due to impact with
other satellites or payloads or natural objects; and debris of satellites
or payloads that are physically attacked or cease to operate due to
military action.

[0027] As space debris is not under active control, it tends to be
unstable. In certain instances, the motion of space debris is a
combination of tumbling, yawing and rotation which, when coupled in three
dimensions, produces unstable, random, or chaotic (i.e., non-uniform)
motion. In the case of larger pieces of space debris, the object may have
large moments of inertia and the unstable motion may generate significant
rotational momentum.

[0028] As used herein, "pneumatic impingement" means impinging an object
with a gas plume. In some embodiments, the gas plume has sufficient force
to generate a torque on the space debris. In some embodiments, the object
is unstable space debris.

[0029] As used herein, "stable" means the state in which dynamic motion of
an object in space, measured relative to a fixed frame of reference, is
reduced to about zero (or, is substantially zero) about at least two of
the three principal axes of motion and the dynamic motion around the
third is not accelerating, or is decelerating to zero.

[0030] As used herein, "stabilize" means the application of force to an
object in space, that is in unconstrained motion, so as to reduce the
dynamic motion of that object, measured relative to a fixed reference
frame, to about zero (or, to substantially zero) about at least two of
the three principal axes of motion and the dynamic motion around the
third is not accelerating, or is decelerating to about zero.

[0031] As used herein, "stabilized" means space debris with reduced
dynamic motion, measured relative to a fixed reference frame. In some
embodiments, the dynamic motion has been reduced to about zero (or
substantially zero) about at least two of the three principal axes of
motion and the dynamic motion around the third is not accelerating, or is
decelerating to zero

[0032] As used herein, the terms "gas plume", "gas jet" and "gas stream"
are used interchangeably and mean a body of gas that is expelled through
an opening at high speed. The gas moves in the same direction at
(generally) the same time. None of the terms imply the body of gas has a
particular shape. In some embodiments, the gas plume is divergent, linear
(i.e., non-divergent), or convergent (i.e., focused).

[0033] As used herein, "orbital parameters" means the three spatial
dimensions which define position, the velocity in each of these
dimensions, as measured in reference to an inertial frame of reference,
and the acceleration in each of these dimensions, as measured in
reference to an inertial frame of reference.

[0034] As used herein, "keep out space" means the minimum separation
between unstable space debris and a satellite disclosed herein to ensure
the safety of the satellite. In some embodiments, the keep out space is
defined by the volume of space in which unstable space debris rotates,
yaws and/or tumbles.

[0035] As used herein, "station keeping operations" means maneuvers used
to keep a spacecraft in an assigned orbit or at a specified distance and
direction from another object in space.

[0036] As used herein, the phrase "the rotational motion is substantially
zero" means that the rotational motion is zero, is about 1% of the
original rotational motion, about 2% of the original rotational motion,
about 3% of the original rotational motion, about 4% of the original
rotational motion, about 5% of the original rotational motion, about 6%
of the original rotational motion, about 7% of the original rotational
motion, about 8% of the original rotational motion, about 9% of the
original rotational motion, or about 10% of the original rotational
motion.

Current Approaches to Space Debris Stabilization

[0037] Space debris includes inoperative satellites and payloads, expended
rocket stages which are not de-orbited, and the residual debris from of
satellites and payloads which catastrophically fail. While functioning
normally, satellites and payloads are actively or passively controlled by
internal control systems. At the "end of life" (EOL) of these systems,
those satellites and payloads that are not de-orbited are uncontrolled
and therefore may become unstable. Additionally, satellites and payloads
may unexpectedly fail before the designed EOL at which time they may
become uncontrollable and unstable. As a result they may begin to tumble
or rotate in one or more axes of motion (becoming unstable). Expended
rocket stages which are not de-orbited, and residual debris from of
satellites and payloads which catastrophically fail, may have no control
system and their rotational motion is may also be unstable. There is a
need for methods of stabilizing unstable space debris.

Methods of Stabilizing Unstable Space Debris

[0038] Disclosed herein, in certain embodiments, are methods of
stabilizing unstable space debris. In some embodiments, the methods
comprise applying a force to precise points on the debris (i.e., the
target points).

[0039] In some embodiments, the methods comprise pneumatic impingement at
precise points on the debris. In some embodiments, the methods comprise
impinging specific target point(s) on the debris with a gas plume [See
FIG. 2]. In some embodiments, each target point is impinged
simultaneously. In some embodiments, each target point is impinged
sequentially.

[0040] In some embodiments, a gas plume is applied with sufficient force
to dampen all moments of inertia of the debris. In some embodiments, a
gas plume is applied until the rotational motion of space debris is
reduced to about zero (or, the rotational motion is substantially zero)
about at least two of the three axes of rotation.

[0041] In some embodiments, a pulse of a gas plume is applied to each
target point. As used herein, "pulse" means a single and abrupt emission
of a gas plume. The number of pulses depends on the total moment of
inertia on each of the axes of rotation and the amount of force that each
pulse imparts on each target point.

[0042] In some embodiments, a single pulse is applied to each target
point. In some embodiments, a series of pulses (e.g., short pulses) is
applied to each target point. In some embodiments, the number of pulses
of a gas plume required to stabilize an axis of motion on the unstable
space debris is independent of the number of pulses of a gas plume
required to dampen the moments of inertia about the other axes of motion
of the unstable space debris. In some embodiments, the change of the
rotational momentum about an axis is calculated after each pulse. In some
embodiments, the application of the gas to a target point is halted when
the rotational motion about the axis of motion has stabilized. In some
embodiments, the application of the gas to a target point is halted when
the rotational motion is reduced to about zero (or, the rotational motion
is substantially zero) about at least two of the three axes of rotation.

[0043] In order to stabilize unstable space debris that is unstable in all
three axes, there must be at least one "target point" on, or near to,
each of the mutually orthogonal axes. In order to access each of the
target points, the opening through which a gas plume is expelled must be
oriented such that the gas plume will impact the target points but is
maintained outside the keep-out space. In some embodiments, the opening
through which a gas jet is expelled is located on a maneuverable arm
which is maneuvered into place. In some embodiments, the opening through
which a gas jet is expelled is fixed to the satellite and maneuvered into
place via positioning of the satellite and maintenance of the position
and attitude while the pneumatic impingement system is operating.

Characteristics of the Gas Used for Pneumatic Impingement

[0044] Disclosed herein, in certain embodiments, are methods of
stabilizing unstable space debris. In some embodiments, the methods
comprise pneumatic impingement of the unstable space debris at specific
target point(s). In some embodiments, the gas jet comprises any suitable
gas. Factors influencing the suitability of a gas include, but are not
limited to, whether the gas is accelerated via pressure (compression),
whether the gas is accelerated via vaporization, whether the gas is
accelerated via combustion, whether the gas is accelerated via exothermic
chemical decomposition of a monopropellant reacting with a catalyst,
whether the gas is accelerated via electromagnetic means, toxicity, cost,
density, easy of handling, ease of storage, or a combination thereof.

[0045] In some embodiments, the methods comprise impinging target point(s)
on the unstable space debris with a plume of cold or warm gas. As used
herein, "cold gas" means a gas that has been pressurized by means of
compression or evaporation and that is accelerated by the pressure
differential between the internal pressures of the pressure vessel that
the pressurized gas is contained in and the external pressure of the
ambient environment ("space") and then expelled through an opening. The
opening is any suitable opening. In some embodiments, the opening is a
nozzle. In some embodiments, the opening is a divergent type nozzle, a
convergent type nozzle, a linear-type (or, collimated-type) nozzle, or a
combination thereof.

[0046] As used herein, "warm gas" means a gas that is accelerated and then
expelled through an opening by means of vaporization. As used herein,
"vaporization" means heating the liquid gas past its boiling point so
that it transitions into a gas. The pressure in the chamber in which the
liquid is vaporized increases significantly as the liquid vaporizes. The
gas is accelerated as it is expelled through an opening by the pressure
differential between the internal pressures in the vaporization chamber
and the external pressure of the ambient environment ("space"). Exemplary
gases that can be accelerated and then expelled through an opening by (a)
compression or (b) vaporization include, but are not limited to: hydrogen
(H2), helium (He2), xenon (Xe), argon (Ar), neon (Ne), freon,
gaseous ammonia, or nitrogen (N2). The opening is any suitable
opening. In some embodiments, the opening is a nozzle. In some
embodiments, the opening is a divergent type nozzle, a convergent type
nozzle, a linear-type (or, collimated-type) nozzle, or a combination
thereof.

[0047] In some embodiments, the methods comprise impinging target point(s)
on the unstable space debris with a plume of hot gas. In some
embodiments, the hot gas is the high velocity residual affluent resulting
from chemical combustion of an oxidizer (e.g., liquid oxygen (LOX),
gaseous oxygen (GOX), fluorine (F2), oxygen difluoride (OF2),
Tetrafluorohydrazine (N2F2), Chlorine pentafluoride (CIF5)
or other similar oxidizers) and a propellant (e.g., hydrogen (H2),
Kerosene products such as Rocket Propellant 1 (RP-1) and Rocket
Propellant 2 (RP-2), Methane (CH4), Monomethylhydrazine (MMH) or
other such propellants) which is expelled at high velocity through an
opening. In some embodiments, the gas is the high velocity residual
affluent from the chemical decomposition of a monopropellant such as (but
limited to) Hydrazine (N2H4) or hydrogen peroxide
(H2O2) caused by an exothermic reaction between the propellant
and a catalyst such as (but not limited to) iridium, silver or platinum
which is expelled at high velocity through an opening.

[0048] In some embodiments, the methods comprise impinging target point(s)
on the unstable space debris with a gas plume. In some embodiments, the
gas is accelerated by electrothermal, electrostatic, or electromagnetic
means which is expelled at high velocity through an opening. Exemplary
gases that can be accelerated by electrothermal, electrostatic, or
electromagnetic means include, but are not limited to, hydrogen
(H2), helium (He2), xenon (Xe), argon (Ar), neon (Ne), nitrogen
(N2), ionized cesium (Cs), or ionized mercury (Mg).

[0049] In some embodiments, the opening through which the gas is expelled
is situated relative to the unstable space debris such that the gas will
impinge on target point(s) on the unstable space debris at the proper
angle of impact.

[0050] In some embodiments, the opening through which the gas jet is
expelled is any suitable size. The suitable size of the opening is
dependent on multiple factors including, but not limited to, engineering
analysis, how much force is to be generated, how long the pulses of gas
should last, the distance from the opening to the point the gas jet is
focused on, the temperature of the gas jet, and back pressure.

[0051] In some embodiments, the opening though with the accelerated gas is
expelled is a divergent type nozzle which forms a conically shaped gas
plume that diverges following expulsion from the nozzle. The portion of
the affluent which impinges on the target points when the gas plume is
formed by a divergent type nozzle is a function of the rate of
divergence, the size of the target points, the distance from the nozzle,
or any combination thereof; and will decrease proportionally as the range
from the nozzle to the target point increases. In some embodiments, the
opening though with the accelerated gas is expelled is a linear type (or,
collimated-type) nozzle which forms a collimated gas plume that neither
substantially diverges nor converges for a specified distance from the
nozzle. The portion of the affluent which impinges on the target points
when a linear nozzle forms a collimated gas plume is a function of the
size of the collimated gas plume, the size of the target points, the
pointing accuracy of the system used to direct the gas plume at the
targets points, or any combination thereof; and will remain relatively
constant with range (until the gas plume begins to diverge naturally in
the far field). In some embodiments, the opening though with the
accelerated gas is expelled is a convergent type nozzle which forms a gas
plume that substantially converges the affluent to a specific point at a
specified range (distance) from the nozzle following which it begins to
diverge. The portion of the affluent which impinges on the target points
when the gas plume is formed by a convergent nozzle is a function of the
rate of convergence, the size of the target points, pointing accuracy of
the system used to direct the gas plume at the targets points, the range
from the nozzle, or any combination thereof; and will increase
proportionally until the gas plume reaches the focus point at which point
it will decrease as the range from the nozzle increases.

Determination of Target Points

[0052] Disclosed herein, in certain embodiments, are methods of
stabilizing the unstable space debris. In some embodiments, the methods
comprise application of force to precise target points on the unstable
space debris such that the force will generate a torque on the unstable
space debris that will dampen or counter the rotational momentum about
one or more of the axes of rotation. In some embodiments, the methods
comprise application of force to precise target points on the unstable
space debris such that the force will reduce rotational motion to about
zero (or, the rotational motion is substantially zero) about at least two
of the three axes of rotation. In some embodiments, the methods comprise
pneumatic impingement of the debris at precise target points.

[0053] In some embodiments, the target points on the unstable space debris
are determined by analysis of the motion of the unstable space debris. In
some embodiments, the methods comprise identifying (a) the Center of Mass
and (b) the direction and magnitude of the velocity vector of the Center
of Mass of the body of the unstable space debris. In some embodiments,
the methods comprise identifying three mutually orthogonal axes of motion
centered at the Center of Mass. In some embodiments, the methods comprise
identifying the rotational direction, rates, moments of inertia, and
rotational momentum of the body of the unstable space debris relative to
the three mutually orthogonal axes of motion centered at the Center of
Mass. In some embodiments, the target points are structural locations on
the unstable space debris body. In some embodiments, the target points
are structural locations on the unstable space debris body that are on,
or near to, one of the three mutually orthogonal axes of motion. In some
embodiments, the target points are structural locations on the unstable
space debris body that are structurally rigid enough to absorb the force
that the gas plume imparts on the unstable space debris. In some
embodiments, the target points are structural locations on the unstable
space debris body that are (a) on, or near to, one of the three mutually
orthogonal axes of motion, and (b) structurally rigid enough to absorb
the force that the gas plume imparts.

Determination of the Amount of Applied Force

[0054] Disclosed herein, in certain embodiments, are methods of
stabilizing unstable space debris. In some embodiments, the methods
comprise the application of force that generates a torque (e.g.,
pneumatic impingement) to the unstable space debris. In some embodiments,
the force is applied at precise target points. In some embodiments, the
force is applied until the rotational motion of the space debris is
reduced to about zero (or, the rotational motion is substantially zero)
about at least two of the three axes of rotation.

[0055] In some embodiments, the amount of force applied to the unstable
space debris depends on (a) the motion of the debris, (b) the state
vector for the Center of Mass of the unstable space debris, (c) the
mutually orthogonal axes, (d) the Moments of Inertia, (e) rotational
momentum and (f) the target points. In some embodiments, the methods
comprise calculating the force to be applied to the debris.

[0056] In some embodiments, the force applied to the unstable space debris
does not exceed the structural limitations of the target points. In some
embodiments, the force applied to the unstable space debris does not
result in structural failure. In some embodiments, the force applied to
the unstable space debris does not result in the break up the unstable
space debris.

[0057] In some embodiments, the structural limitations are determined by
an analysis of the object (e.g., by analyzing the design schematics of
the debris and the materials used to construct the debris). In some
embodiments, the force applied to a target point is calculated before
impingement (e.g., following a preceding impingement).

[0058] In some embodiments, the structural limitations are determined by a
visual inspection of the debris, scanning of the debris, application of
minimal force to the debris with a gradual increase in the force vector
as necessary, or a combination thereof.

[0059] In some embodiments, the force applied to a target point is a
function of the velocity of the gas plume and the mass of the gas plume
that impacts the target point.

Modularity

[0060] In some embodiments, the space debris is large space debris. In
some embodiments, large space debris that is in unstable random
rotational motion defines a large volume of keep-out space. In certain
instances, it is impractical to use deployable arms that are long enough
to avoid the keep-out space generated by large space debris. Thus, in
some embodiments, a method disclosed herein uses multiple satellites. In
some embodiments, a method disclosed herein uses one satellite per axis
of rotation. In some embodiments, a method disclosed herein uses multiple
satellites per axis of rotation.

[0061] In some embodiments, each satellite is positioned such that (a) it
has access to a target point, and (b) it avoids the keep-out space.

[0062] In some embodiments, each satellite operates independently. For
example, each satellite is independently positioned via an earth-based
operator.

[0063] In some embodiments, the satellites coordinate with each other
autonomously. In some embodiments, the satellites coordinate with each
other autonomously and are monitored via an earth-based operator.

[0064] In some embodiments, each satellite is capable of independently
stabilizing small or medium sized space debris.

Sources of Force Other than Pneumatic Impingement

[0065] There are multiple methods that can be used to apply force to
unstable space debris. In some embodiments, the force is applied via
physical contact with a mechanical element (e.g., an arm or part of the
satellite bus). However, the energy stored in the inertia motion or
rotational momentum of the system may be sufficiently large to result in
the structural failure of the debris or the mechanical element, thus
creating additional debris.

[0066] In some embodiments, the force is applied via contacting the
unstable space debris with a viscous liquid (e.g., water or a more
complex liquid). In some embodiments, the viscosity and mass of the
liquid imparts a force on the unstable space debris thus slowing its
rotation.

[0067] In some embodiments, the force is applied via illuminating the
unstable space debris with a laser. In some embodiments, the laser
illuminates one side of the unstable space debris which vaporizes some of
the material. In some embodiments, as the material is vaporized and
ejected from the unstable space debris, it creates a force in the equal
and opposite direction that is a function of the mass of the material
being vaporized and the velocity at which it is expelled. In some
embodiments, the laser creates a radiation pressure differential between
the side of the space debris that the laser illuminates and the side of
the space debris that the laser does not illuminate and this radiation
pressure differential is sufficient to dampen the rotational momentum
about one of more axes of rotation.

[0068] In some embodiments, the force is applied by generating an
electromagnetic field that the unstable space debris rotates within. In
some embodiments, the unstable space debris rotating in an
electromagnetic field creates electrical currents (i.e., eddy currents)
on the conductive material that the space debris is constructed of. In
some embodiments, the interaction of the electrical currents and
naturally occurring planetary magnetic fields creates an electromagnetic
force on the unstable space debris which dampens the rotation of the
object.

[0069] In some embodiments, the force is applied by contacting the
unstable space debris with particulate material (e.g., sand or some other
material). In some embodiments, the impact of the particulate material
imparts a force and creates a drag about the axes of motion and slows the
unstable space debris.

Methods of Altering the Orbit of Space Debris or Capturing Space Debris

[0071] Disclosed herein, in certain embodiments, is a method of capturing
space debris. In some embodiments, the space debris is small space debris
with low rotational energy states. In some embodiments, the space debris
is unstable and it is stabilized prior to capture. In some embodiments,
the unstable space debris is stabilized by a method disclosed herein.

[0072] In some embodiments, a satellite disclosed herein and the space
debris rendezvous. In some embodiments, the satellite matches energy
states (e.g., orbital velocity) with the space debris. In some
embodiments, the satellite captures the space debris by any suitable
means. In some embodiments, the satellite captures the space debris by
use of a deployable arm, bag, tow line, net, magnetic affecter,
mechanical affecter or a combination thereof.

Changing Orbital Parameters

[0073] Disclosed herein, in certain embodiments, are methods of altering
the orbital path of space debris. In some embodiments, the methods
comprise changing the orbital parameters of space debris. In some
embodiments, the space debris is unstable and it is stabilized prior to
capture. In some embodiments, the methods first comprise stabilizing
unstable space debris by a method disclosed here.

[0074] In some embodiments, the orbital parameters of the space debris are
modified via attachment of a de-orbit module to the debris. In some
embodiments, the de-orbit module is attached to the unstable space debris
via use of a robotic arm. In some embodiments, the de-orbit module is
attached to the unstable space debris by mechanically transferring the
rocket pack following docking.

[0075] In some embodiments, the orbital parameters of the space debris are
modified via application of force at precise points of the debris (i.e.,
the target points) intended to change the linear momentum and velocity
vector of the space debris. In some embodiments, the orbital parameters
of space debris are modified via pneumatic impingement at precise points
on the debris.

[0076] In some embodiments, the methods comprise applying a force to
specific target point(s) on the debris. In some embodiments, the methods
comprise impingement by a gas plume at specific target point(s) on the
debris [See FIG. 2]. In some embodiments, the gas plume impacts the
debris with sufficient force to change the linear momentum and direction
and magnitude of the velocity vector of the space debris.

[0077] In some embodiments, pulses of gas plumes are applied to one or
more target points. In some embodiments, the gas plumes are applied to
the target point(s) for an extended period of time. In some embodiments,
the duration of each gas plume depends on the linear momentum and the
total change of the direction and magnitude of the velocity vector of the
unstable space debris that is required.

[0078] In some embodiments, the required velocity vector change is
achieved by a single gas plume impingement event. In some embodiments,
the required velocity vector change is achieved by multiple gas plume
impingement events. In some embodiments, change in the velocity vector of
the space debris is determined (e.g., calculated) after each gas
impingement event. In some embodiments, the gas impingement on the target
point(s) is halted when the required change in direction and magnitude of
the unstable space debris velocity vector is achieved.

Satellite

[0079] Disclosed herein, in certain embodiments, are satellites for
stabilizing an unstable space debris, comprising: (a) satellite bus with
standard subsystems and interfaces, (b) a means for generating and
projecting a force sufficient to dampen the rotational momentum about one
or more of the axes of rotation of the unstable space debris, (c) one or
more external sensors (radar, laser radar (LIDAR), optical, or imaging
sensors), and (d) and an electronic system designed to analyze the data
from the sensors and develop a stabilization plan which includes the
force, duration, number, direction and magnitude of the pneumatic (gas)
jets (plumes).

[0080] In some embodiments, a satellite disclosed herein is further
capable of capturing the space debris which has been stabilized. In some
embodiments, a satellite disclosed herein is further capable of altering
the orbital path of space debris that has been stabilized.

[0081] In some embodiments, a satellite disclosed herein applies force to
the space debris (e.g., unstable space debris, or space debris that has
been stabilized). In some embodiments, a satellite disclosed herein
applies force to unstable space debris to stabilize the debris. In some
embodiments, a satellite disclosed herein applies force to the stabilized
space debris to change the orbital path of the space debris. In some
embodiments, a satellite disclosed herein applies force to the unstable
space debris by pneumatic impingement. In some embodiments, pneumatic
impingement results from a gas plume directed from the satellite.

Design

[0082] In some embodiments, the satellite comprises: a means for
generating a force sufficient to dampen the rotational momentum about one
or more of the axes of rotation of the unstable space debris or to alter
the linear motion of the satellite and therefore changing its orbital
parameters.

[0083] In some embodiments, the means for generating the force is a means
for compressing a gas. A gas is compressed when it is forced, at high
pressure, by a pump into a container or tank (a.k.a., pressure vessel). A
gas may also be compressed by the injection into a pressure vessel
containing an unpressurized gas of a benign (non-reactive) high pressure
gas or liquid. The gas or liquid that is injected into the pressure
vessel to compress the operating gas may be injected directly into the
pressure vessel or it may be isolated from the operating gas that is
being compressed by a bladder inside the pressure vessel designed to
separate the operating gas and the compressing gas or liquid. The
pressure vessel, having sufficient strength to hold the gas at high
pressure, is connected by a system of tubes, connectors, and valves to an
opening. The opening is any suitable opening. In some embodiments, the
opening is part of a nozzle. In some embodiments, the opening is a
divergent type nozzle, linear type (or, collimated-type) nozzle, a
convergent type nozzle, or a combination thereof. When the valves are
opened in the correct sequence, the gas is accelerated by the pressure
differential between the internal pressures of the pressure vessel that
the pressurized gas is contained in and the external pressure of the
ambient environment ("space") and then expelled through the opening. The
acceleration is a function of the pressure differential and the design of
the opening. The force that is generated is a function of the
acceleration and the mass of the gas that is expelled. The mass of the
gas is a function of the pressure and the length of time (duration) that
the values are open and closed. Exemplary gases that can be accelerated
and then expelled through an opening by compression include, but are not
limited to: hydrogen (H2), helium (He2), xenon (Xe), argon
(Ar), neon (Ne), freon, gaseous ammonia, or nitrogen (N2).

[0084] In some embodiments, the means for generating the force is a means
for vaporizing a liquefied gas. "Vaporization" means heating the liquid
gas past its boiling point so that it transitions into a gas. The
pressure in the chamber in which the liquid is vaporized increases
significantly as the liquid vaporizes. The gas is accelerated as it is
expelled through the opening by the pressure differential between the
internal pressures in the vaporization chamber and the external pressure
of the ambient environment ("space"). Exemplary gases that can be
accelerated and then expelled through an opening by vaporization include,
but are not limited to: hydrogen (H2), helium (He2), xenon
(Xe), argon (Ar), neon (Ne), freon, gaseous ammonia, or nitrogen
(N2).

[0085] In some embodiments, the means for generating the force is a means
for combustion of an oxidizer and a propellant. The force is generated by
the acceleration to a high velocity of the residual affluent resulting
from chemical combustion of an oxidizer (e.g., liquid oxygen (LOX),
gaseous oxygen (GOX), fluorine (F2), oxygen difluoride (OF2),
Tetrafluorohydrazine (N2F2), Chlorine pentafluoride (CIF5)
or other similar oxidizers) and a propellant (e.g., hydrogen (H2),
Kerosene products such as Rocket Propellant 1 (RP-1) and Rocket
Propellant 2 (RP-2), Methane (CH4), Monomethylhydrazine (MMH) or other
such propellants) after injection in a combustion chamber. After the
oxidizer and propellant are injected and mixed in the combustion chamber,
they can be ignited by the discharge of an electrical spark or current,
explosive charge, laser heating or other methods. The combustion of the
explosive mix of oxidizer and propellant become self sustaining and
caused the pressure and temperature in the combustion chamber is
significantly increased creating a highly energized affluent. The force
that is generated is a function of the acceleration and the mass of the
affluent that is expelled. The affluent is accelerated as it is expelled
through the throat and then the expansion section of a nozzle by the
pressure differential between the internal pressures in the combustion
chamber and the external pressure of the ambient environment ("space").

[0086] In some embodiments, the means for generating the force is through
the exothermic decomposition of a monopropellant. The force is generated
by the acceleration to a high velocity of the residual affluent resulting
from the chemical decomposition of a monopropellant such as (but limited
to) Hydrazine (N2H4) or hydrogen peroxide (H2O2)
caused by an exothermic reaction between the propellant and a catalyst
such as (but not limited to) iridium, silver or platinum. Valves are
opened or closed to control the amount of monopropellant is injected into
the combustion chamber through a catalyst, the monopropellant decomposes
exothermically, converting from a liquid to a high pressure and
temperature gas. The force that is generated is a function of the
acceleration, the mass of the affluent that is expelled and the duration
that the valves are open so that the monopropellant can be injected into
the combustion chamber. The affluent is accelerated as it is expelled
through the throat and then the expansion section of a nozzle by the
pressure differential between the internal pressures in the combustion
chamber and the external pressure of the ambient environment ("space").

[0087] In some embodiments, the means for generating the force is by an
electromagnetic means. In some embodiments, the force is generated when a
gas is energized and accelerated by electrothermal, electrostatic, or
electromagnetic means. As used herein, "electrothermal" means
electromagnetic devices where electromagnetic fields are used to generate
a plasma to increase the heat of the bulk propellant. In some
embodiments, a plasma is accelerated by the Lorentz force resulting from
the interaction between the current flowing through the plasma and the
magnetic field (which is either externally applied, or induced by the
current) out through the exhaust chamber. The force that is generated is
a function of the acceleration and the mass of the plasma or ion stream
that is expelled. The plasma or ion stream is accelerated as it is
expelled through the throat and then the expansion section of a nozzle.

[0089] In some embodiments, the system which produces energized gas
electrostaticly is an electrostatic ion thruster, a Hall affect thruster,
a field emission electric propulsion system, or any combination thereof.
In each case, ions are accelerated by the potential difference of a
static electrical field between an anode and a cathode.

[0090] In some embodiments, the system which produces energized gas
electromagnetically is an electrodeless plasma thruster, pulsed inductive
thruster, helicon double layer thruster, a Magnetoplasmadynamic (MPD)
thruster, or any combination thereof.

[0091] In some embodiments, the means for generating the force is by
accelerating a liquid or particulate. In some embodiments, a liquid is
compressed when it is forced, at high pressure, by a pump into a
container or tank (a.k.a., pressure vessel). In some embodiments, a
liquid is compressed by the injection directly into a pressure vessel
containing an unpressurized liquid of a benign (non-reactive) high
pressure gas or liquid. The gas or liquid that is injected into the
pressure vessel to compress the operating liquid may be injected directly
into the pressure vessel or it may be isolated from the operating liquid
that is being compressed by a bladder inside the pressure vessel designed
to separate the operating liquid and the compressing gas or liquid. The
pressure vessel, having sufficient strength to hold the gas at high
pressure, is connected by a system of tubes, connectors, and valves to an
opening. The opening is any suitable opening. In some embodiments, the
opening is part of a nozzle. The pressure vessel, having sufficient
strength to hold the liquid at high pressure, is connected by a system of
tubes, connectors, and valves to the opening. When the valves are opened
in the correct sequence, the liquid is accelerated by the pressure
differential between the internal pressures of the pressure vessel that
the pressurized liquid is contained in and the external pressure of the
ambient environment ("space") and then expelled through the opening. The
acceleration is a function of the pressure differential and the design of
the opening. The force that is generated is a function of the
acceleration and the mass of the liquid that is expelled. The mass of the
liquid is a function of the pressure and the length of time (duration)
that the values are open and closed. In some embodiments, a particulate
is used. In some embodiments, the particulate is injected into a high
pressure gas pulse similar to the compressed gas system described above.
The use of a particulate increases the mass and kinetic energy of the
compressed gas jet plume.

[0092] In some embodiments, the means for generating the force is a laser.
The laser illuminates one side of the unstable space debris which
vaporizes some of the material. As the material is vaporized and ejected
from the unstable space debris, it creates a force in the equal and
opposite direction that is a function of the mass of the material being
vaporized and the velocity at which it is expelled. Alternatively, the
laser can produce a radiation pressure differential between the side of
the space debris that the laser illuminates and the side of the space
debris that the laser does not illuminate and the radiation pressure
differential is sufficient to dampen the rotational momentum about one of
more axes of rotation.

[0093] In some embodiments, the means for generating the force is a means
for generating an electromagnetic field. An electromagnetic field is
created whenever an electrical current is generated between two points.
In some embodiments, an electrical current flowing in an antenna will
create an electromagnetic field that emanates from the antenna. The shape
of the antenna will determine the shape and directionality of the
electromagnetic field. An electrical current will be generated in a
conductive material that moves within an electromagnetic field. In some
embodiments, the satellite using an electrical power source generates an
electrical current in an antenna which will create an electromagnetic
field that encompasses (shape and directionality) the space debris. Since
the space debris is rotating within the electromagnetic field, and it may
be constructed of so electrically conductive material, the rotational
motion of the debris within the electromagnetic field will generate small
electrical currents on, or within, the debris. The electrical currents
on, or within, the debris will simultaneously interact with the naturally
occurring planetary magnetic fields which surround the Earth creating a
force that will attempt to align itself with those naturally occurring
planetary magnetic fields. As the electrical currents on, or within the
debris, begin to align themselves with the naturally occurring planetary
magnetic fields they will generate a force that will tends to dampen the
rotational motion of the debris in one or more of the axes of rotation.

[0094] In some embodiments, a satellite disclosed herein comprises a
carriage for the means for generating a force sufficient to dampen the
rotational momentum about one or more of the axes of rotation of the
unstable space debris (e.g., a gas, a liquid, particulate matter, a
laser, an electromagnetic field). In some embodiments, the carriage
components comprise a storage tank, container, or pressure vessel. In
some embodiments, the carriage components comprise a means for moving the
source of the force (e.g., a gas, a liquid, particulate matter) from the
storage tanks, containers or pressure vessels to the opening. In some
embodiments, the means for moving the source of the force from the
storage compartment to the opening is a series of interconnected pipes.

[0095] In some embodiments, a satellite disclosed herein comprises a
carriage for the pneumatic components. In some embodiments, the carriage
of the pneumatic components comprises a storage compartment. In some
embodiments, the carriage of the pneumatic components comprises a means
for moving the gas from the storage compartment to the opening. In some
embodiments, the means for moving the gas from the storage compartment to
the opening is a series of interconnected pipes.

[0096] In some embodiments, the satellite comprises an opening through
which the force moves. The opening is any suitable opening. In some
embodiments, the opening is part of a nozzle. In some embodiments, the
opening is a divergent type nozzle, linear type nozzle, a convergent type
nozzle, or a combination thereof. In some embodiments, the opening though
with the accelerated gas is expelled is a divergent type nozzle which
forms a conically shaped gas plume that diverges following expulsion from
the nozzle. The portion of the affluent which impinges on the target
points when the gas plume is formed by a divergent type nozzle is a
function of the rate of divergence, the size of the target points, the
distance from the nozzle, or any combination thereof; and will decrease
proportionally as the range from the nozzle to the target point
increases. In some embodiments, the opening though with the accelerated
gas is expelled is a linear type nozzle which forms a collimated gas
plume that neither substantially diverges nor converges for a specified
distance from the nozzle. The portion of the affluent which impinges on
the target points when a linear nozzle forms a collimated gas plume is a
function of the size of the collimated gas plume, the size of the target
points, the pointing accuracy of the system used to direct the gas plume
at the targets points, or any combination thereof; and will remain
relatively constant with range (until the gas plume begins to diverge
naturally in the far field). In some embodiments, the opening though with
the accelerated gas is expelled is a convergent type nozzle which forms a
gas plume that substantially converges the affluent to a specific point
at a specified range (distance) from the nozzle following which it begins
to diverge. The portion of the affluent which impinges on the target
points when the gas plume is formed by a convergent nozzle is a function
of the rate of convergence, the size of the target points, pointing
accuracy of the system used to direct the gas plume at the targets
points, the range from the nozzle, or any combination thereof; and will
increase proportionally until the gas plume reaches the focus point at
which point it will decrease as the range from the nozzle increases.

[0097] In some embodiments, the opening through which the gas jet is
expelled is any suitable size. In some embodiments, the suitable size of
the opening is a function of engineering analysis, how much force is to
be generated, the planned duration of the gas plume, the distance from
the opening to the point that the gas jet is focused on, the temperature
of the gas jet, back pressure, or any combination thereof.

[0098] In some embodiments, the opening is located on the main satellite
body. In some embodiments, the opening is located on an arm attached to
the satellite.

[0099] In some embodiments, a satellite disclosed herein comprises the
physical structure (i.e., satellite bus) that contains normal satellite
subsystems and external internal faces (e.g., physical structure, the
computers, the wires, the batteries, the GNC system, the navigation
sensors, the environmental control systems, the propulsion system, the
communications system) As used herein, "satellite bus" means all the
elements of a satellite described herein except for the payload
(comprised of the pneumatic (plume) impingement system, the scanning
sensors, de-orbit module, means for collecting the debris, and means for
storing the debris).

[0100] In some embodiments, a satellite disclosed herein comprises a
guidance system, navigation and control (GNC) system (including
stabilization and attitude control), or a combination thereof.

[0101] In some embodiments, a satellite disclosed herein comprises a means
for managing satellite subsystems and external interfaces. In some
embodiments, the means for managing satellite subsystems and external
interfaces is through a computer based vehicle management system (VMS).
In some embodiments, the VMS operates autonomously, is operated remotely
from a ground control station, or a combination of both.

[0102] In some embodiments, a satellite disclosed herein comprises a means
for powering the satellite. In some embodiments, a satellite disclosed
herein comprises a power management and distribution system. In some
embodiments, the means for powering the satellite is any suitable
powering mechanism. Examples of powering mechanisms include, but are not
limited to: solar panels, thermal batteries, fuel cells, nuclear
reactors, or a combination thereof.

[0103] In some embodiments, a satellite disclosed herein comprises a means
for propulsion. In some embodiments, the means for propulsion is any
suitable propulsion mechanism. Examples of propulsion mechanisms include,
but are not limited to: chemical rocket engines, nuclear rocket engines,
cold gas rocket engines, or electrical rocket engines.

[0104] In some embodiments, a satellite disclosed herein comprises a means
for communication. In some embodiments, a satellite disclosed herein
comprises multiple means for communication. In some embodiments, the
means for communication comprises at least one antenna for receiving
communications and transmitting communications. In some embodiments,
communications are received and/or transmitted via radio waves,
microwaves, optical frequency, or a combination thereof. In some
embodiments, the radio is a physical radio. In some embodiments, the
radio is a software radio. In some embodiments, the means for
communication further comprises a computer module capable of encrypting
communications.

[0107] In some embodiments, the satellite further comprises a de-orbit
module for attachment to the space debris. In some embodiments, the
satellite comprises a means for attaching the de-orbit module to the
unstable space debris. In some embodiments, the de-orbit module is
attached to the unstable space debris via use of a robotic arm or by
mechanical latches. In some embodiments, the de-orbit module is attached
to the unstable space debris by mechanically transferring the rocket pack
following docking.

[0108] In some embodiments, the satellite further comprises a means for
collecting the unstable space debris. In some embodiments, the means for
collecting the debris is a robotic arm. In some embodiments, the means
for collecting the debris is a bag that scoops up the debris. In some
embodiments, the means for collecting the debris is a net. In some
embodiments, the means for collecting the debris is a magnetic affecter
that magnetically attracts and then latches on to the ferrous metal
elements of the unstable space debris.

[0109] In some embodiments, the satellite further comprises a means for
storing the unstable space debris. In some embodiments, the means for
storing the unstable space debris is a bag attached to the satellite. In
some embodiments, the means for storing the unstable space debris is a
box attached to the satellite. In some embodiments, the means for storing
the unstable space debris is a tow line attached to the satellite. In
some embodiments, the space debris is attached to the tow line via
magnetic attraction, use of a hook, use of a clamp, use of a harpoon, or
a combination thereof.

Dimensions

[0110] In some embodiments, the total volume of gas required to stabilize
unstable space debris is directly proportional to the total moment of
inertia of the unstable space debris.

[0111] In certain instances, the volume of gas or liquid required to
stabilize a large space debris is significantly larger than the volume of
gas or liquid required to stabilize a small space debris. Consequently,
in some embodiments, the satellite is sized to stabilize the largest body
of space debris that is anticipated. Alternatively, in some embodiments,
the satellite is scalable and is produced in various overall sizes.

Maneuverability

[0112] In some embodiments, a satellite disclosed herein rendezvous with
the unstable space debris. In some embodiment, a satellite disclosed
herein maneuvers to rendezvous with the unstable space debris by changing
its orbit to parallel the velocity vector of the Center of Mass of the
unstable space debris. In some embodiments, a satellite disclosed herein
maneuvers itself such that (a) it has access to the three mutually
orthogonal axes (e.g., with one or more deployable arms or with one or
satellites), and (b) it maintains a position outside the "keep out space"
volume.

[0113] In some embodiments, a satellite disclosed herein conducts
proximity operations by maneuvering around the unstable space debris to
(a) survey the debris (e.g., determine the condition of the debris), and
(b) identify potential target points. In some embodiments, a satellite
disclosed herein will then conduct station keeping operations to maintain
position while the motion of the unstable space debris is analyzed. [See
FIG. 4].

[0114] In some embodiments, the expulsion of a gas by the satellite
imparts a force (e.g., linear (i.e., thrust) or torque (i.e., rotation))
on the satellite. In some embodiments, the force propels or rotates the
satellite away from the unstable space debris. In some embodiments, a
satellite disclosed herein maintains its orientation referenced to the
mutually orthogonal X, Y and Z axes of the unstable space debris. In some
embodiments, a satellite disclosed herein has an active control system.
In some embodiments, the active control system comprises a reaction
control system (RCS), control moment gyroscopes (CMG), magnetic torque
converters for attitude control, or a combination thereof. In some
embodiments, the active control system is sized to perform station
keeping and to maintain attitude sufficient to counter the forces
imparted on the satellite by the gas impingement system as it is
operating. In some embodiments, a satellite disclosed herein further
comprises a main engine and an orbital maneuvering system or thruster.

Arms

[0115] In certain instances, in order to stabilize the unstable space
debris that is unstable in all three axes, there must be at least one
"target point" on, or near to, each of the mutually orthogonal axes. In
order to access each of the target points, the opening through which a
gas plume is expelled must be oriented such that the gas plume will
impact the target points but is maintained outside the keep-out space. In
some embodiments, the opening through which a gas jet is expelled is
maneuvered into place by use of at least one maneuverable arm. In some
embodiments, a satellite disclosed herein comprises at least two arms. In
some embodiments, a satellite disclosed herein comprises three arms.

[0116] In some embodiments, each arm is mechanically articulated. In some
embodiments, each mechanically articulated arm is flexible instead of
rigid.

Motion Analysis

[0117] In some embodiments, the satellite analyzes the motion of the
unstable space debris. In some embodiments, the motion of the debris is
analyzed via use of a laser tracking system. In some embodiments, the
satellite comprises a means for laser tracking. The laser tracking system
(or Laser Radar--LIDAR) consists of a laser transmitter that can transmit
short, low power laser pulses at the space debris and a receiver that can
detect the reflected laser pulses and measure the time (of arrival) and
direction of the reflected laser pulse. The data from the laser tracking
system is processed by a computer to develop a three dimensional virtual
model of the debris which can be used for analysis to determine
structural integrity and locations and orientation of the space debris
and its components. The laser tracking system can also track specific
points on the space debris and identify the velocity, direction and
acceleration of those points. That data from the laser tracking system is
computer analyzed to determine the rotational axes, rotational rates,
rotational momentum, and rotational direction of the space debris as part
of developing and executing the stabilization plan.

[0118] In some embodiments, the satellite analyzes the motion of the
unstable space debris. In some embodiments, the motion of the debris is
analyzed via use of a radar (or other radio frequency) tracking system.
In some embodiments, a satellite comprises a means for radar tracking.
The radar (or other radio frequency) tracking system consists of a
microwave frequency transmitter that can transmit short, low power
electromagnetic pulses at the space debris and a receiver that can detect
the reflected electromagnetic pulses and measure the time (of arrival)
and direction of the reflected pulse. The data from the radar tracking
system is processed by a computer to develop a three dimensional virtual
model of the debris which can be used for analysis to determine
structural integrity and locations and orientation of the space debris
and its components. The radar tracking system can also track specific
points on the space debris and identify the velocity, direction and
acceleration of those points. That data from the radar tracing system is
computer analyzed to determine the rotational axes, rotational rates,
rotational momentum, and rotational direction of the space debris as part
of developing and executing the stabilization plan.

[0119] In some embodiments, the satellite analyzes the motion of the
unstable space debris. In some embodiments, the motion of the debris is
analyzed via use of an optical tracking system. In some embodiments, a
satellite comprises a means for optical tracking. An optical tracking
system consists of one of more optical sensor(s) (receivers) that
collects ambient reflected light (defined as a passive optical system) or
reflected light generated by a light source (defined as an active optical
systems) (transmitter). A single optical tracking sensor can track color
differences, light and dark images, edges, read symbols and wording, and
can determine the direction and two dimensional motion of specific target
points. An optical tracking system consisting of more than one sensor can
additionally use binocular vision to determine the range to specific
target points and determine three dimensional motion. The data from the
optical tracking system is processed by a computer to develop a three
dimensional virtual model of the debris which can be used for analysis to
determine structural integrity and the location and orientation of the
space debris and its components. The optical tracking system can also
track specific points on the space debris and identify the velocity,
direction and acceleration of those points. That data from the optical
tracking system is computer analyzed to determine the rotational axes,
rotational rates, rotational momentum, and rotational direction of the
space debris as part of developing and executing the stabilization plan.

[0120] In some embodiments, the satellite analyzes the motion of the
unstable space debris. In some embodiments, the motion of the debris is
analyzed via use of a combination of (a) laser tracking systems, (b)
radar (or other radio frequency) tracking systems, and/or (c) optical
tracing systems. In some embodiments, the satellite comprises a means for
(a) laser tracking, (b) radar (or other radio frequency) tracking, and/or
(c) optical tracking.

Calculations

[0121] In some embodiments, a satellite disclosed herein comprises a means
for calculating the amount of force to be applied to the unstable space
debris, the number and timing of the gas pulses to be applied to the
unstable space debris, or a combination thereof. In some embodiments, the
means for performing the aforementioned calculations is an on-board
computer module. In some embodiments, the means for performing these
calculations is a ground-based computer module that communicates (as
previously described) with the satellite. In some embodiments, the means
for performing these calculations is a combination of on-board and ground
based systems.

Reusability

[0122] Most space vehicles are built to be expendable at "end of life"
(EOL) due to the difficulty of refueling and remotely maintaining a space
vehicle on orbit. In some embodiments, the satellite disclosed herein is
reusable. In some embodiments, a satellite disclosed herein is
maintainable (i.e., components can be removed and replaced). In some
embodiments, a satellite disclosed herein is refuelable (both maneuvering
propellant and impingement gas).

[0123] While preferred embodiments of the present invention have been
shown and described herein, it will be obvious to those skilled in the
art that such embodiments are provided by way of example only. Numerous
variations, changes, and substitutions will now occur to those skilled in
the art without departing from the invention. It should be understood
that various alternatives to the embodiments of the invention described
herein may be employed in practicing the invention. It is intended that
the following claims define the scope of the invention and that methods
and structures within the scope of these claims and their equivalents be
covered thereby.

EXAMPLES

Example 1

Method of Stabilizing Unstable Space Debris Via Pneumatic Impingement

[0124] There are several hundred non-functioning rocket bodies (r/b) and
spacecraft (s/c) in Low Earth Orbit (LEO) between 600 kilometers (km) and
2,000 km altitude which create the potential for an impact that will
increase the population of space debris in LEO. In order to reduce the
probability of collision, a decision is made to capture and de-orbit
several of these rocket bodies and spacecraft each year. After surveying
the non-functioning spacecraft and rocket bodies in that altitude band,
it is determined that Rocket Body One has the highest potential for
collision and so it becomes the highest priority to de-orbit. It is also
determined that the attitude and orientation of Rocket Body One is
unstable, that it exhibits characteristics of rotational motion in three
axes, and that the rotational motion is coupled between all three axes so
that the rotational motion appears to be random. Rocket Body One must be
stabilized before it can be captured and de-orbited. Pneumatic (plume)
impingement is used to stabilize the rocket body.

[0125] The pneumatic (plume) impingement stabilization spacecraft executes
a rendezvous with Rocket Body One. This is done by using the main
propulsion system to propel the pneumatic (plume) impingement
stabilization spacecraft from its initial orbit into the orbit of Rocket
Body One. As the pneumatic (plume) impingement stabilization spacecraft
begins to rendezvous with Rocket Body One, sensors onboard the pneumatic
(plume) impingement stabilization spacecraft track Rocket Body One to
measure its dimensions. As Rocket Body One rotates, it sweeps out a
spherical volume of space about its center of mass--the "keep out space".
The pneumatic (plume) impingement stabilization spacecraft remains
outside this "keep out space" by maintaining at least a range from the
center of mass of Rocket Body One that is greater than the radius of the
spherical volume of space that Rocket Body One sweeps out as it rotates
about its center of mass.

[0126] Next, the pneumatic (plume) impingement orbital debris
stabilization spacecraft maneuvers around Rocket Body One and conducts a
scan of Rocket Body One determining the structural integrity and
developing a three dimensional map of Rocket Body One. The information,
developed from the scan of Rocket Body One by sensors onboard the
pneumatic (plume) impingement stabilization spacecraft, is used as an
input by the pneumatic (plume) impingement system planner to develop a
stabilization plan that will not exceed structural integrity limits based
on external appendages, such as antennae and solar panels, or other
residual structural limitations identified during the scan.

[0127] Once the condition of the rocket body is determined, the linear
motion is characterized, the axes about which Rocket Body One is rotating
are determined, and the rotational motion of Rocket Body One is
determined by analysis of the data measured during the survey of Rocket
Body One. The pneumatic (plume) impingement system planner develops a
debris stabilization plan based on this data. The debris stabilization
plan consists of (at least) one target point on an axis of motion
perpendicular to each of the axes about which the debris is rotating.
Since Rocket Body One is rotating about all three axes of rotation, the
debris stabilization plan requires three targets points. The planner also
develops a sequence of thruster firings designed to create a sequence of
pneumatic jets (plumes) that will impinge on the target points,
transferring the force of those pneumatic jets (plumes) to the target
points which creates torques on Rocket Body One that are opposite to the
rotations of Rocket Body One. The total sequence of thruster firings is
designed to reduce the rotational motion of Rocket Body One to zero about
at least two of the three axes of rotation. The planner also determines
pulse duration and variations in the force of each thruster firing. The
force transferred to each target point is a function of the force of each
pneumatic jet (plume) and the angle of impact of the plume on the target
point. Since Rocket Body One is rotating, the time and angle each target
point is exposed to the pneumatic jet (plume) varies as a function of the
rotational rate of Rocket Body One. The planner also develops a plan for
the pneumatic (plume) impingement orbital debris stabilization spacecraft
guidance, navigation and control system which will counter the forces
imparted on the pneumatic (plume) impingement orbital debris
stabilization spacecraft by the firing of the pneumatic (plume)
impingement, so that the pneumatic (plume) impingement orbital debris
stabilization spacecraft can perform station keeping in order to maintain
its orientation and distance to Rocket Body One.

[0128] The pneumatic (plume) impingement orbital debris stabilization
spacecraft positions itself and orients the pneumatic (plume)
stabilization system nozzles to aim at the target points based on the
stabilization plan. The pneumatic impingement system generates pulses of
pneumatic jets (gas plumes) consisting of the affluent generated by the
combustion of a fuel and oxidizer carried by the pneumatic (plume)
impingement orbital debris stabilization spacecraft (bi-propellant base
system).

[0129] These pulses are directed at the selected target points on Rocket
Body One. Each pulse is timed to impact the target points while the
target points are perpendicular to the pneumatic jets (gas plumes) and
the pneumatic jets (gas plumes) are tangential to the arc of rotation.
The pneumatic (plume) impingement orbital debris stabilization spacecraft
uses its guidance, navigation and control system to maintain its position
and orientation by countering the force that the pneumatic impingement
system imparted on the pneumatic impingement orbital debris stabilization
spacecraft.

[0130] While the pneumatic (plume) impingement orbital debris
stabilization spacecraft is executing the pneumatic impingement plan, it
tracks the rocket body with sensors to determine if the torque that is
generated by the impingement of the gas plume on the target points is
reducing the rotational energy and motion of the rocket body as planned.
The pneumatic (plume) impingement system planner processes the data from
the sensors and uses that feedback to continuously monitor and update the
stabilization plan.

[0131] The pneumatic (plume) impingement orbital debris stabilization
spacecraft continues to execute the stabilization plan, monitoring the
affect on Rocket Body One, processes the feedback and updates the
stabilization plan, until the rotational motion is reduced to zero in the
selected axes of rotation.

[0132] There are several hundred non-functioning rocket bodies (r/b) and
spacecraft (s/c) in Low Earth Orbit (LEO) between 600 kilometers (km) and
2,000 km altitude which create the potential for an impact that will
increase the population of space debris in LEO. In order to reduce the
probability of collision, a decision is made to capture and de-orbit
several of these rocket bodies and spacecraft each year. After surveying
the non-functioning spacecraft and rocket bodies in that altitude band,
it is determined that Rocket Body Two has a high potential for collision
and so it becomes a high priority to de-orbit. It is also determined that
the attitude and orientation of Rocket Body Two is unstable, that it
exhibits characteristics of rotational motion in three axes, and that the
rotational motion is coupled between all three axes so that the
rotational motion appears to be random. Rocket Body Two must be
stabilized and then captured in order to be de-orbited. Pneumatic (plume)
impingement is used to stabilize Rocket Body Two. The pneumatic (plume)
impingement stabilization spacecraft will then capture Rocket Body Two
and affix a mechanical device to de-orbit it.

[0133] The pneumatic (plume) impingement stabilization spacecraft executes
a rendezvous with Rocket Body Two. This is done by using the main
propulsion system to propel the pneumatic (plume) impingement
stabilization spacecraft from its initial orbit into the orbit of Rocket
Body Two. As the pneumatic (plume) impingement stabilization spacecraft
begins to rendezvous with Rocket Body Two, sensors onboard the pneumatic
(plume) impingement stabilization spacecraft track Rocket Body Two to
measure its dimensions. As Rocket Body Two rotates, it sweeps out a
spherical volume of space about its center of mass--the "keep out space".
The pneumatic (plume) impingement stabilization spacecraft remains
outside this "keep out space" by maintaining at least a range from the
center of mass of Rocket Body Two that is greater than the radius of the
spherical volume of space that Rocket Body Two sweeps out as it rotates
about its center of mass.

[0134] Next, the pneumatic (plume) impingement orbital debris
stabilization spacecraft maneuvers around Rocket Body Two and conducts a
scan of Rocket Body Two determining the structural integrity, developing
a three dimensional map of Rocket Body Two, and identifying structurally
rigid points on Rocket Body Two where Rocket Body Two can be mechanically
captured. The information, developed from the scan of Rocket Body Two by
sensors onboard the pneumatic (plume) impingement stabilization
spacecraft, is used as an input by the pneumatic (plume) impingement
system planner to develop a stabilization plan that will not exceed
structural integrity limits based on external appendages, such as
antennae and solar panels, or other residual structural limitations
identified during the scan.

[0135] Once the condition of the rocket body is determined, the linear
motion is characterized, the axes about which Rocket Body Two is rotating
are determined, and the rotational motion of Rocket Body Two is
determined by analysis of the data measured during the survey of Rocket
Body Two. The pneumatic (plume) impingement system planner develops a
debris stabilization plan based on this data. The debris stabilization
plan consists of (at least) one target point on an axis of motion
perpendicular to each of the axes about which the debris is rotating.
Since Rocket Body Two is rotating about all three axes of rotation, the
debris stabilization plan requires three targets points. The planner also
develops a sequence of thruster firings designed to create a sequence of
pneumatic jets (plumes) that will impinge on the target points,
transferring the force of those pneumatic jets (plumes) to the target
points which creates torques on Rocket Body Two that are opposite to the
rotations of Rocket Body Two. The total sequence of thruster firings is
designed to reduce the rotational motion of Rocket Body Two to zero about
at least two of the three axes of rotation. The planner also determines
pulse duration and variations in the force of each thruster firing. The
force transferred to each target point is a function of the force of each
pneumatic jet (plume) and the angle of impact of the plume on the target
point. Since Rocket Body Two is rotating, the time and angle each target
point is exposed to the pneumatic jet (plume) varies as a function of the
rotational rate of Rocket Body Two. The planner also develops a plan for
the pneumatic (plume) impingement orbital debris stabilization spacecraft
guidance, navigation and control system which will counter the forces
imparted on the pneumatic (plume) impingement orbital debris
stabilization spacecraft by the firing of the pneumatic (plume)
impingement, so that the pneumatic (plume) impingement orbital debris
stabilization spacecraft can perform station keeping in order to maintain
its orientation and distance to Rocket Body Two.

[0136] The pneumatic (plume) impingement orbital debris stabilization
spacecraft positions itself and orients the pneumatic (plume)
stabilization system nozzles to aim at the target points based on the
stabilization plan. The pneumatic impingement system generates pulses of
pneumatic jets (gas plumes) consisting of the affluent generated by the
combustion of a fuel and oxidizer carried by the pneumatic (plume)
impingement orbital debris stabilization spacecraft (bi-propellant base
system).

[0137] These pulses are directed at the selected target points on Rocket
Body Two. Each pulse is timed to impact the target points while the
target points are perpendicular to the pneumatic jets (gas plumes) and
the pneumatic jets (gas plumes) are tangential to the arc of rotation.
The pneumatic (plume) impingement orbital debris stabilization spacecraft
uses its guidance, navigation and control system to maintain its position
and orientation by countering the force that the pneumatic impingement
system imparted on the pneumatic impingement orbital debris stabilization
spacecraft.

[0138] While the pneumatic (plume) impingement orbital debris
stabilization spacecraft is executing the pneumatic impingement plan, it
tracks the rocket body with sensors to determine if the torque that is
generated by the impingement of the gas plume on the target points is
reducing the rotational energy and motion of the rocket body as planned.
The pneumatic (plume) impingement system planner processes the data from
the sensors and uses that feedback to continuously monitor and update the
stabilization plan.

[0139] The pneumatic (plume) impingement orbital debris stabilization
spacecraft continues to execute the stabilization plan, monitoring the
affect on rocket body two, processes the feedback and updates the
stabilization plan, until the rotational motion is reduced to zero in the
selected axes of rotation.

[0140] Once the pneumatic (plume) impingement orbital debris stabilization
spacecraft stabilizes the rotational motion of Rocket Body Two by
eliminating the rotational motion in at least two of the three axes of
motion, it proceeds to capture Rocket Body Two by mechanically latching
onto the main propulsion system thrust nozzle. The pneumatic (plume)
impingement orbital debris stabilization spacecraft maneuvers to a
position in alignment with the thrust nozzle and with sufficient
separation to ensure that the end effecters are clear of the thrust
nozzle. Using the pneumatic (plume) impingement orbital debris
stabilization spacecraft thrusters, the stabilization spacecraft slowly
closes the range between the stabilization spacecraft and rocket body two
maintaining alignment with the rocket body two thrust nozzle. The
pneumatic (plume) impingement orbital debris stabilization spacecraft
thrusters also match any rocket body two residual rotation rates that are
not eliminated during the stabilization process. Closure rates are
controlled such that the impact forces will not damage either spacecraft,
or will cause a reaction bounce that will force a separation between the
spacecraft. Four articulated arms with articulated end effectors are
positioned so that one end effector is outside the thrust cone and one
end effector is inside the thrust cone. After contact, the end effecters
are engaged to capture rocket body two.

[0141] Once the capture is completed, the pneumatic (plume) impingement
orbital debris stabilization spacecraft releases a de-orbit module and
moves to separate itself and Rocket Body Two. Once there is sufficient
between the stabilization spacecraft and rocket body two, the de-orbit
module is activated and it imparts sufficient force on rocket body two to
create a linear deceleration of Rocket Body Two that will cause it to
lose altitude and ultimately de-orbit.

[0142] There are several hundred non-functioning rocket bodies (r/b) and
spacecraft (s/c) in Low Earth Orbit (LEO) between 600 kilometers (km) and
2,000 km altitude which create the potential for an impact that will
increase the population of space debris in LEO. In order to reduce the
probability of collision, a decision is made to capture and de-orbit
several of these rocket bodies and spacecraft each year. After surveying
the non-functioning spacecraft and rocket bodies in that altitude band,
it is determined that Spacecraft One has a high potential for collision
and so it becomes a high priority to de-orbit. It is also determined that
the attitude and orientation of Spacecraft One is unstable, that it
exhibits characteristics of rotational motion in three axes, and that the
rotational motion is coupled between all three axes so that the
rotational motion appears to be random. Spacecraft One must be stabilized
and then captured in order to be de-orbited. Pneumatic (plume)
impingement is used to stabilize Spacecraft One. The pneumatic (plume)
impingement stabilization spacecraft will then alter the orbit of
Spacecraft One to de-orbit it.

[0143] The pneumatic (plume) impingement stabilization spacecraft executes
a rendezvous with Spacecraft One. This is done by using the main
propulsion system to propel the pneumatic (plume) impingement
stabilization spacecraft from its initial orbit into the orbit of
Spacecraft One. As the pneumatic (plume) impingement stabilization
spacecraft begins to rendezvous with Spacecraft One, sensors onboard the
pneumatic (plume) impingement stabilization spacecraft track Spacecraft
One to measure its dimensions. As Spacecraft One rotates, it sweeps out a
spherical volume of space about its center of mass--the "keep out space".
The pneumatic (plume) impingement stabilization spacecraft remains
outside this "keep out space" by maintaining at least a range from the
center of mass of Spacecraft One that is greater than the radius of the
spherical volume of space that Spacecraft One sweeps out as it rotates
about its center of mass).

[0144] Next, the pneumatic (plume) impingement orbital debris
stabilization spacecraft maneuvers around Spacecraft One and conducts a
scan of Spacecraft One determining the structural integrity and
developing a three dimensional map of Spacecraft One. The information,
developed from the scan of Spacecraft One by sensors onboard the
pneumatic (plume) impingement stabilization spacecraft, is used as an
input by the pneumatic (plume) impingement system planner to develop a
stabilization plan that will not exceed structural integrity limits based
on external appendages, such as antennae and solar panels, or other
residual structural limitations identified during the scan.

[0145] Once the condition of the rocket body is determined, the linear
motion is characterized, the axes about which Spacecraft One is rotating
are determined, and the rotational motion of Spacecraft One is determined
by analysis of the data measured during the survey of Spacecraft One. The
pneumatic (plume) impingement system planner develops a debris
stabilization plan based on this data. The debris stabilization plan
consists of (at least) one target point on an axis of motion
perpendicular to each of the axes about which the debris is rotating.
Since Spacecraft One is rotating about all three axes of rotation, the
debris stabilization plan requires three targets points. The planner also
develops a sequence of thruster firings designed to create a sequence of
pneumatic jets (plumes) that will impinge on the target points,
transferring the force of those pneumatic jets (plumes) to the target
points which creates torques on Spacecraft One that are opposite to the
rotations of Spacecraft One. The total sequence of thruster firings is
designed to reduce the rotational motion of Spacecraft One to zero (or,
to substantially zero) about at least two of the three axes of rotation.
The planner also determines pulse duration and variations in the force of
each thruster firing. The force transferred to each target point is a
function of the force of each pneumatic jet (plume) and the angle of
impact of the plume on the target point. Since Spacecraft One is
rotating, the time and angle each target point is exposed to the
pneumatic jet (plume) varies as a function of the rotational rate of
Spacecraft One. The planner also develops a plan for the pneumatic
(plume) impingement orbital debris stabilization spacecraft guidance,
navigation and control system which will counter the forces imparted on
the pneumatic (plume) impingement orbital debris stabilization spacecraft
by the firing of the pneumatic (plume) impingement, so that the pneumatic
(plume) impingement orbital debris stabilization spacecraft can perform
station keeping in order to maintain its orientation and distance to
Spacecraft One. In addition, since the pneumatic (plume) impingement
orbital debris stabilization spacecraft will use its pneumatic (plume)
stabilization system to alter the orbital path of Spacecraft One, the
stabilization plan is developed so that Spacecraft One, once it is
stabilized, is oriented with a structural element (in this case, the main
propulsion system thrust nozzle) that is aligned with the velocity vector
of Spacecraft One and is on the leading side of Spacecraft.

[0146] The pneumatic (plume) impingement orbital debris stabilization
spacecraft positions itself and orients the pneumatic (plume)
stabilization system nozzles to aim at the target points based on the
stabilization plan. The pneumatic impingement system generates pulses of
pneumatic jets (gas plumes) consisting of the affluent generated by the
combustion of a fuel and oxidizer carried by the pneumatic (plume)
impingement orbital debris stabilization spacecraft (bi-propellant base
system).

[0147] These pulses are directed at the selected target points on
Spacecraft One. Each pulse is timed to impact the target points while the
target points are perpendicular to the pneumatic jets (gas plumes) and
the pneumatic jets (gas plumes) are tangential to the arc of rotation.
The pneumatic (plume) impingement orbital debris stabilization spacecraft
uses its guidance, navigation and control system to maintain its position
and orientation by countering the force that the pneumatic impingement
system imparted on the pneumatic impingement orbital debris stabilization
spacecraft.

[0148] While the pneumatic (plume) impingement orbital debris
stabilization spacecraft is executing the pneumatic impingement plan, it
tracks the rocket body with sensors to determine if the torque that is
generated by the impingement of the gas plume on the target points is
reducing the rotational energy and motion of the rocket body as planned.
The pneumatic (plume) impingement system planner processes the data from
the sensors and uses that feedback to continuously monitor and update the
stabilization plan.

[0149] The pneumatic (plume) impingement orbital debris stabilization
spacecraft continues to execute the stabilization plan, monitoring the
affect on Spacecraft One, processes the feedback and updates the
stabilization plan, until the rotational motion is reduced to zero in the
selected axes of rotation.

[0150] The stabilization plan is executed such that Spacecraft One is
oriented with the main propulsion system thrust nozzle aligned with the
velocity vector of Spacecraft One. The main propulsion system thrust
nozzle is the target point that the pneumatic (plume) impingement orbital
debris stabilization spacecraft will use to impart a force, using
pneumatic impingement, opposite to the direction of Spacecraft One's
velocity vector--this force being sufficient to reduce the magnitude of
the velocity vector of Spacecraft One so that Spacecraft One will
de-orbit.

[0151] The pneumatic (plume) impingement orbital debris stabilization
spacecraft maneuvers to a position in alignment with the velocity vector
of Spacecraft One. Using the pneumatic (plume) impingement orbital debris
stabilization spacecraft thrusters, the stabilization spacecraft fires a
continuous gas jet (plume) targeted at Spacecraft One's mail propulsion
system thrust nozzle. The gas jet (plume) cuts off, once the velocity of
Spacecraft One is decelerated sufficiently to ensure de-orbit of
Spacecraft One. During the firing of the stabilization spacecraft's
thrusters, the pneumatic (plume) impingement orbital debris stabilization
spacecraft uses its internal GNC and RCS systems to maintain its position
relative to Spacecraft One. After Spacecraft One's orbital parameters are
altered sufficiently to de-orbit Spacecraft One, the pneumatic (plume)
impingement orbital debris stabilization spacecraft's internal GNC and
RCS systems are used to stabilize pneumatic (plume) impingement orbital
debris stabilization spacecraft's orbit so that it does not de-orbit with
Spacecraft One.

[0152] There are several hundred non-functioning rocket bodies (r/b) and
spacecraft (s/c) in Low Earth Orbit (LEO) between 600 kilometers (km) and
2,000 km altitude which create the potential for an impact that will
increase the population of space debris in LEO. In order to reduce the
probability of collision, a decision is made to capture and de-orbit
several of these rocket bodies and spacecraft each year. After surveying
the non-functioning spacecraft and rocket bodies in that altitude band,
it is determined that Rocket Body Three has a high potential for
collision and so it becomes a high priority to de-orbit. It is also
determined that the attitude and orientation of Rocket Body Three is
unstable, that it exhibits characteristics of rotational motion in three
axes, and that the rotational motion is coupled between all three axes so
that the rotational motion appears to be random. Rocket Body Three must
be stabilized and then captured in order to be de-orbited. An orbital
debris stabilization spacecraft will use an electromagnetic field to
stabilize Rocket Body Three. The orbital debris stabilization spacecraft
will then capture Rocket Body Three and affix a mechanical device to
de-orbit it.

[0153] The orbital debris stabilization spacecraft executes a rendezvous
with Rocket Body Three. This is done by using the main propulsion system
to propel the orbital debris stabilization spacecraft from its initial
orbit into the orbit of Rocket Body Three. As the orbital debris
stabilization spacecraft begins to rendezvous with Rocket Body Three,
sensors onboard the orbital debris stabilization spacecraft track Rocket
Body Three to measure its dimensions. As Rocket Body Three rotates, it
sweeps out a spherical volume of space about its center of mass--the
"keep out space". The orbital debris stabilization spacecraft remains
outside this "keep out space" by maintaining at least a range from the
center of mass of Rocket Body Three that is greater than the radius of
the spherical volume of space that Rocket Body Three sweeps out as it
rotates about its center of mass.

[0154] Next, the orbital debris stabilization spacecraft maneuvers around
Rocket Body Three and conducts a scan of Rocket Body Three determining
the structural integrity of Rocket Body Three to determine if it has
sufficient structural integrity to be subject to electromagnetic
stabilization, if there is sufficient conductive material to support
electromagnetic stabilization and to determined the rotational motion of
Rocket Body Three by analysis of the data measured during the survey of
Rocket Body Three.

[0155] The orbital debris stabilization spacecraft positions itself and
orients its antenna toward Rocket Body Three so that Rocket Body Three
will rotate within the electromagnetic field generated by the orbital
debris stabilization spacecraft.

[0156] The orbital debris stabilization spacecraft powers the
electromagnetic field orbital debris stabilization system, generating an
electromagnetic field that encompasses Rocket Body Three and which
generates electrical currents ("eddy currents") on, or within, the
conductive material which is part of Rocket Body Three. As Rocket Body
Three continues to rotate within the electromagnetic field being created
by the orbital debris stabilization spacecraft, the electrical currents
("eddy currents") on, or within, the conductive material which is part of
Rocket Body Three interact with the naturally occurring planetary
magnetic fields. The forces created by this interaction attempt to align
the electrical currents ("eddy currents") on, or within, the conductive
material which is part of Rocket Body Three and the naturally occurring
planetary magnetic fields. As the rotation of Rocket Body Three begins to
dampen out by the force between the eddy currents and the naturally
occurring planetary magnetic field, the electrical currents ("eddy
currents") on, or within, the conductive material which is part of Rocket
Body Three tend to weaken. Ultimately the rotational rates about at least
two of the three axes are reduced to zero (or, to substantially zero) and
Rocket Body Three is stabilized.

[0157] Once the orbital debris stabilization spacecraft stabilizes the
rotational motion of Rocket Body Three by eliminating the rotational
motion in at least two of the three axes of motion, it proceeds to
capture Rocket Body Three by mechanically latching onto the main
propulsion system thrust nozzle. The orbital debris stabilization
spacecraft maneuvers to a position in alignment with the thrust nozzle
and with sufficient separation to ensure that the end effecters are clear
of the thrust nozzle. Using the orbital debris stabilization spacecraft
thrusters, the stabilization spacecraft slowly closes the range between
the orbital debris stabilization spacecraft and Rocket Body Three
maintaining alignment with the Rocket Body Three thrust nozzle. The
orbital debris stabilization spacecraft thrusters also match any Rocket
Body Three residual rotation rates that are not eliminated during the
stabilization process. Closure rates are controlled such that the impact
forces will not damage either spacecraft, or will cause a reaction bounce
that will force a separation between the spacecraft. Four articulated
arms with articulated end effectors are positioned so that one end
effector is outside the thrust cone and one end effector is inside the
thrust cone. After contact, the end effecters are engaged to capture
Rocket Body Three.

[0158] Once the capture is completed, the orbital debris stabilization
spacecraft releases a de-orbit module and moves to separate itself and
Rocket Body Three. Once there is sufficient between the orbital debris
stabilization spacecraft and Rocket Body Three, the de-orbit module is
activated and it imparts sufficient force on Rocket Body Three to create
a linear deceleration of Rocket Body Three that will cause it to lose
altitude and ultimately de-orbit.

Example 5

Testing and Verification through Computer Simulation

[0159] Testing and verification of the pneumatic (plume) impingement
orbital debris stabilization spacecraft is achieved through computer
simulation prior to implementation in an operational system. This same
type of computer simulation is also usable as part of the planning
process for actual missions to verify the stabilization plan prior to
implementation and to verify that the expected affects of the plume
impingement process match the simulated plan.

[0160] A computer simulation will require a virtual model of the pneumatic
(plume) impingement orbital debris stabilization spacecraft, a physics
based algorithm of the method of imparting force on the unstable space
debris, a virtual model of the unstable space debris (including
specification of rotational motion and orbital parameters), and a
simulation of the space environment (including gravitation forces and
other external forces and pressures that the unstable space debris and
the pneumatic (plume) impingement orbital debris stabilization spacecraft
will be subject to).

[0161] The computer simulation is used test and verify the full system
performance, the performance of each subsystem, simulate the motion of
the unstable space debris, simulate the affect of imparting forces on the
unstable space debris, and all other mission phases.

Example 6

Testing and Verification Through Physical Simulation

[0162] Testing and verification of the pneumatic (plume) impingement
orbital debris stabilization spacecraft is achieved through live (or,
physical) methods to verify the suitability and performance of computer
(or, virtual) simulations. Live (or, physical) testing is done on the
ground. Live (or, physical) testing is done in two dimensions and three
dimensions.

[0163] Two dimensional live (or, physical) testing is done on large "air
bearing floors" available at NASA. An "air bearing floor" simulates a
zero resistance two dimensional surface that simulates the motion of an
object in space in two dimensions (x and y axis in a Cartesian coordinate
system) by mounting physical models of the unstable space debris and the
pneumatic (plume) impingement orbital debris stabilization spacecraft on
flat bottomed sleds that then ride on a thin cushion of air that is
injected from the floor of the test facility. Using this system, we test
and verify system performance in two dimensions and replicate roll, pitch
and yaw. After verifying system performance in two dimensions, we extend
the simulation into three dimensions mathematically.

[0164] Three dimensional live (or physical) testing is done using a
similar testing facility with the addition of an "overhead gantry" with a
mechanical arm that can dynamically simulate motion relative to a fixed
object on the ground, or a movable object on an "air Bearing floor". This
type of live simulated environment can replicate 8 degrees of freedom
including bridge, trolley, waist, shoulder, extension, roll, yaw, and
pitch allowing for complete testing and development of this method of
stabilizing unstable space debris.